The present invention relates generally to a turbine cooling structure in a gas turbine engine and more specifically to an improved configuration of a combustor/turbine successive dual cooling arrangement.
In a conventional gas turbine engine comprising a compressor, combustor and turbine, both the combustor and the turbine require cooling due to beating thereof by hot combustion gases.
Within the combustor, fuel fed through the fuel nozzle is mixed with compressed air provided by the high pressure compressor and ignited to drive turbines with the hot gases emitted through the combustor. Within the metal combustor, the gases burn at approximately 3,500xc2x0 to 4,000xc2x0 Fahrenheit. The combustion chamber is fabricated of a metal which can resist extremely high temperatures. However, even highly resistant metal will melt at approximately 2,100xc2x0 to 2,200xc2x0 Fahrenheit. Therefore, it is important to adequately cool the hot combustor wall of a gas turbine engine for safe engine operation.
As is well known in the art, the combustion gases are prevented from directly contacting the material of the combustor through use of a cool air film which is directed along the internal surfaces of the combustor. The combustor has a number of louver openings through which compressed air is fed parallel to the hot combustor walls. Eventually the cool air curtain degrades and is mixed with the combustion gases. However, in such air film cooling arrangements, the cooling air mixed with the combustion gases increases CO emissions. Thus, while cooling techniques used on the combustor liner may be advantageous in increasing maximum engine temperature, they deleteriously increase CO formation and emission.
The use of air film cooling is limited by the amount of air available exclusively for cooling the combustor wall. Generally, as the amount of cooling air is increased to cool the engine components, the amount of air available for the combustor is decreased, which results in increasing NOx formation and emission.
Efforts have been made to cool the combustor wall of a gas turbine engine while avoiding the increase of emissions. For example, U.S. Pat. No. 5,687,572, issued to Schrantz et al. on Nov. 18, 1997, discloses a combustor for a gas turbine engine having a porous outer metallic shell and a thin-walled, nonporous ceramic liner the backside of which is impingement cooled. All air flow used for impingement cooling is re-injected into the combustion process itself, preferably, primarily in the dilution zone of the combustion process so that there is no loss of pressurized air flow from a thermodynamic standpoint, which is advantageous to reduce NOx formation, and also no film cooling on the interior surface of the ceramic liner is introduced to induce CO formation.
In another example, U.S. Pat. No. 5,758,504 issued to Abreu et al. on Jun. 2, 1998 discloses a combustor construction including an interior liner having a plurality of angled holes extending therethrough, arranged in a pre-established pattern defining a centroid, and an exterior liner having a plurality of holes extending therethrough at about 90 degrees. At least a portion of the holes in the exterior liner are radially aligned with the centroid of the holes in the interior to reduce the use of cooling air flow per unit length of the combustor wall, thereby resulting in reduction of CO emissions.
In addition to the combustor cooling, in a turbine section of a gas turbine engine, both rotating turbine components such as blades, disks and retainers, and stationary turbine components such as vanes, shrouds and frames are directly exposed to high temperature combustion gases discharged from the combustor and routinely require cooling. Cooling of the turbine, especially the rotating components, is critical to the proper function and safe operation of the engine. Failure to adequately cool a turbine disk and its blades, for example, by providing cooling air deficient in supply pressure, volumetric flow rate or temperature margin, may be detrimental to the life and mechanical integrity of the turbine.
Balanced with the need to adequately cool the turbine is the desire for high levels of engine operating efficiency, which translate into lower fuel consumption and lower operating costs. Since turbine cooling air is typically drawn from one or more stages of the compressor and channeled by various means such as pipes, ducts and internal passageways to the desired components, such air is not available to be mixed with the fuel, ignited in the combustor and undergo work extraction in the primary gas flow path of the turbine, total cooling flow bled from the compressor is therefore treated as a parasitic loss in the engine operating cycle, it being desirable to keep such loses to a minimum.
Efforts have been made to minimize compressor bleed and concomitant cycle losses, for example, by attempting to control bleed source or cooling circuit parameters, such as source pressure, pressure drop, flow rate or temperature. One example is disclosed in U.S. Pat. No. 5,555,721 issued to Bourneuf et al. on Sep. 17, 1996. Burneuf et al. describe a turbine cooling supply circuit for a gas turbine engine in which the flow of coolant through the engine is directed to minimize temperature rise prior to discharge into the turbine. In addition to being used for combustion, compressor discharge pressurized air, which is disposed within a combustor casing, is utilized to cool components of the turbine section subject to the hottest combustion gases, namely the stage one nozzle, a stage one shroud and the stage one disk. Additional bleed sources for turbine cooling air include an impeller tip forward bleed flow and impeller tip aft bleed flow which are provided to additionally cool the stage two nozzle and stage two shroud respectively, as well as other turbine components. Bourneuf et al. do not address the cooling of the combustor wall and it would be understood from the drawings attached thereto that a film cooling arrangement is intended to be used.
It has been realized that directing air for cooling, rather than combustion control, limits the degree of combustion emission optimization, and the minimization of the amount of combustor cooling is critical to the design of a state of the art low emission gas turbine combustion system. Therefore, there have been continuous efforts in the industry to develop combustor/turbine cooling apparatus and methods for low emission gas turbine engines.
It is one object of the present invention to provide a low emission gas turbine combustion system using an improved cooling method.
It is another object of the present invention to provide a cooling system for a gas turbine engine to significantly reduce the coolant volume in combustor liner cooling.
It is a further object of the present invention to provide a combustor/turbine successive dual cooling to permit all the air typically used to cool the hot end of the engine downstream of the combustor to be used as combustor cooling as well.
In general terms, a method for cooling a gas turbine engine combustor and turbine section comprises, providing a structure; enabling pressurized cooling air to form air flow impingement on an outer surface of a combustor wall for backside cooling of the combustor wall; directing the air flow immediately upon the impingement thereof along the outer surface of the combustor wall, downstream towards a turbine section for further cooling the combustor wall; and providing an access to exhaust combustor backside cooling air flow for cooling the turbine section.
In accordance with one aspect of the present invention, a cooling apparatus for a gas turbine engine having a combustor comprises a wall adapted to be attached to a combustor wall of a gas turbine engine. The wall is in a spaced apart and substantially parallel relationship with respect to an outer surface of the combustor wall to form an air passage between the wall and the combustor wall for conducting cooling air to cool the outer surface of the combustor wall. Means are provided for introducing the cooling air from a pressurized cooling air source into the air passage to cause impingement thereof on the combustor wall for backside cooling. Means are also provided for discharging the cooling air from the air passage to cool a turbine section downstream of the combustor of the gas turbine engine. The means for introducing cooling air preferably includes perforations in the wall in fluid communication with the air passage and a plenum within a combustor casing so that the cooling air introduced from the plenum through the perforations impinges the outer surface of the combustor wall before being directed downstream through the air passage. The means for discharging the cooling air preferably includes an open downstream end of the air passage to provide an access to exhaust combustor backside cooling air, for hot end cooling so that all, and only the exhaust combustor backside cooling air cools the turbine section.
A gas turbine engine combustor according to a preferred embodiment of the present invention includes a one-piece hot combustor wall defining a combustion chamber. The hot combustor wall includes an inner surface in communication with hot combustion gases flowing towards a turbine section, and an outer surface in contact with cooling air. The gas turbine engine combustor further includes a cold combustor wall fixed to the hot combustor wall. The cold combustor wall is substantially parallel to and disposed at a distance from the hot combustor wall, to form an air passage between the hot and cold walls for directing the cooling air towards the turbine section. A plurality of holes extend through the cold combustor wall in fluid communication with the air passage and a primary plenum within a combustor casing, so that pressurized cooling air in the primary plenum enters the holes to cause an impingement on the outer surface of the hot combustor wall for backside cooling thereof before being directed through the air passage towards the turbine section. The air passage has a closed upstream end and an open downstream end thereof to provide an access to exhaust combustor backside cooling air for cooling the turbine section, whereby all, and only the exhaust combustor backside cooling air is directed to cool turbine components.
It is preferable that a front cold combustor wall is fixed to a front section of the hot combustor wall and a rear cold combustor wall is fixed to a rear section of the hot combustor wall. The front cold combustor wall is fixed at an upstream end thereof by an a spacer, to the front section of the hot combustor wall to form the closed upstream end of the air passage. A downstream end of the front cold combustor wall is supported on a casing structure and the open downstream end of the air passage is adapted to discharge the exhaust combustor backside cooling air into a turbine cavity for cooling turbine components. The rear cold combustor wall is fixed at an upstream end thereof by a spacer, to the rear section of the hot combustor wall, to form the closed upstream end of the air passage. The open downstream end of the air passage formed between the rear cold combustor wall and the rear section of the hot combustor wall is in fluid communication with a secondary plenum providing access to the exhaust combustor backside cooling air for cooling the turbine section.
The combustor/turbine successive dual cooling method and structure according to the present invention advantageously permit compressor air to impingement-backside cool the combustor wall before being diverted to downstream for hot end cooling, and the cooling air is never ingested into the combustion system but is only used to cool the combustor wall and the turbine section so that the amount of combustor backside cooling is significantly reduced, thereby resulting in low emissions of the gas turbine engine.